Low weight large fan gas turbine engine

ABSTRACT

A gas turbine engine according to an example of the present disclosure includes, among other things, a fan configured to deliver airflow to a bypass passage, and a core engine configured to rotate the fan. The core engine includes a high pressure turbine section configured to drive a high pressure compressor section, and a low pressure turbine section configured to drive the fan and a low pressure compressor section. The fan has a fan diameter, Dfan, and the low pressure turbine section has a turbine diameter, Dturb.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a continuation of U.S. patent application Ser. No. 14/428,049, which was filed on Mar. 15, 2015, which is a National Stage Entry of PCT Application No. PCT/US2013/025470, filed on Feb. 10, 2013, which claims priority to U.S. Provisional Application Ser. No. 61/708,288, filed Oct. 1, 2012.

BACKGROUND

A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustion section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section. The compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines.

The high pressure turbine drives the high pressure compressor through an outer shaft to form a high spool, and the low pressure turbine drives the low pressure compressor through an inner shaft to form a low spool. The fan section may also be driven by the inner shaft. A direct drive gas turbine engine includes a fan section driven by the low spool such that the low pressure compressor, low pressure turbine and fan section rotate at a common speed in a common direction.

A speed reduction device such as an epicyclical gear assembly may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section so as to increase the overall propulsive efficiency of the engine. In such engine architectures, a shaft driven by one of the turbine sections provides an input to the epicyclical gear assembly that drives the fan section at a reduced speed such that both the turbine section and the fan section can rotate at closer to optimal speeds.

Although geared architectures have improved propulsive efficiency, turbine engine manufacturers continue to seek further improvements to engine performance including improvements to thermal, transfer and propulsive efficiencies.

SUMMARY

A gas turbine engine according to an example of the present disclosure includes a fan configured to deliver airflow to a bypass passage, and a core engine configured to rotate the fan. The core engine includes a high pressure turbine section configured to drive a high pressure compressor section, and a low pressure turbine section configured to drive the fan and a low pressure compressor section. The fan and the low pressure turbine section are configured to rotate at a common speed and in a common direction. The gas turbine engine has a bypass ratio of greater than about 10. The fan has a fan diameter, Dfan, and the low pressure turbine section has a turbine diameter, Dturb. The fan diameter Dfan and the turbine diameter Dturb have an interdependence represented by a scalable ratio Dturb/Dfan that is between 0.5 and 0.65.

In a further embodiment of any of the foregoing embodiments, the turbine diameter Dturb is defined by an outer case surface of the low pressure turbine section.

In a further embodiment of any of the foregoing embodiments, the fan includes a plurality of fan blades, and the fan diameter Dfan is defined by outer peripheral surfaces of the fan blades.

In a further embodiment of any of the foregoing embodiments, the high pressure turbine section includes two stages.

In a further embodiment of any of the foregoing embodiments, the low pressure compressor section includes a greater number of stages than the high pressure turbine section, and includes fewer stages than the high pressure compressor section.

In a further embodiment of any of the foregoing embodiments, the low pressure turbine section has a greater number of stages than the low pressure compressor section.

In a further embodiment of any of the foregoing embodiments, the fan has a pressure ratio of less than about 1.45.

In a further embodiment of any of the foregoing embodiments, the fan includes fewer than 20 fan blades.

In a further embodiment of any of the foregoing embodiments, the low pressure turbine section has a pressure ratio of greater than about 5.

A gas turbine engine according to an example of the present disclosure includes a fan having fewer than 20 fan blades situated at an inlet of a bypass passage, and a core engine configured to rotate the fan. The core engine includes a high pressure turbine section configured to drive a high pressure compressor section, and a low pressure turbine section configured to drive the fan and a low pressure compressor section. The low pressure turbine section has a greater number of stages than the low pressure compressor section. The fan has a fan diameter, Dfan, the low pressure turbine section has a turbine diameter, Dturb, and the fan diameter Dfan and the turbine diameter Dturb have an interdependence represented by a scalable ratio Dturb/Dfan that is between 0.5 and 0.65.

In a further embodiment of any of the foregoing embodiments, the turbine diameter Dturb is defined by an outer case surface of the low pressure turbine section.

In a further embodiment of any of the foregoing embodiments, the fan diameter Dfan is defined by outer peripheral surfaces of the fan blades.

In a further embodiment of any of the foregoing embodiments, the high pressure turbine section includes two stages.

In a further embodiment of any of the foregoing embodiments, the low pressure compressor section includes a greater number of stages than the high pressure turbine section, and includes fewer stages than the high pressure compressor section.

In a further embodiment of any of the foregoing embodiments, the fan has a pressure ratio of less than about 1.5.

In a further embodiment of any of the foregoing embodiments, the gas turbine engine has a bypass ratio of greater than about 10.

In a further embodiment of any of the foregoing embodiments, the low pressure turbine has a pressure ratio of greater than about 5.

A method of designing a gas turbine engine according to an example of the present disclosure includes providing a fan configured to deliver airflow to a bypass duct, and providing a core engine configured to rotate the fan. The core engine include a high pressure turbine section configured to drive a high pressure compressor section, and a low pressure turbine section configured to rotate the fan at a common speed and in a common direction. The fan has a fan diameter, Dfan, the low pressure turbine section has a diameter, Dturb, and the fan diameter Dfan and the low turbine section diameter Dturb have an interdependence represented by a scalable ratio Dturb/Dfan that is between 0.5 and 0.65. The fan is configured to deliver a portion of air into the core engine, and a portion of air into the bypass duct, and a bypass ratio, which is defined as a volume of air passing to the bypass duct compared to a volume of air passing into the core engine, is greater than 10.

In a further embodiment of any of the foregoing embodiments, the fan has a pressure ratio of less than about 1.5 and the low pressure turbine has a pressure ratio of greater than about 5.

A further embodiment of any of the foregoing embodiments includes a low pressure compressor section driven by the low pressure turbine section. The low pressure compressor section includes a greater number of stages than the high pressure turbine section, and includes fewer stages than the high pressure compressor section, and the high pressure turbine section includes two stages.

In a featured embodiment, a gas turbine engine has a propulsor including a fan and a fan drive geared architecture. The fan defines a fan diameter. A gas generator includes a fan drive turbine, which drives the fan through the fan drive geared architecture. The fan drive turbine has a diameter less than 0.50 the size of the fan diameter.

In another embodiment according to the previous embodiment, the diameter of the fan drive turbine is greater than 0.30 the size of the fan diameter.

In another embodiment according to any of the previous embodiments, the diameter of the fan drive turbine is between about 0.35 and about 0.45 the size of the fan diameter.

In another embodiment according to any of the previous embodiments, the fan drive turbine further comprises a high pressure turbine located upstream of the low pressure turbine.

In another embodiment according to any of the previous embodiments, the fan drive turbine comprises a low pressure turbine.

In another embodiment according to any of the previous embodiments, a compressor section has a low pressure compressor driven by the low pressure turbine and a combustor in fluid communication with the compressor section.

In another embodiment according to any of the previous embodiments, a first shaft connects the low pressure turbine, low pressure compressor, and the fan drive geared architecture.

In another embodiment according to any of the previous embodiments, the fan drive geared architecture comprises an epicyclic gear box.

In another embodiment according to any of the previous embodiments, the diameter of the fan drive turbine is defined by an outer case surface of the fan drive turbine.

In another embodiment according to any of the previous embodiments, the fan diameter is defined by an outer peripheral surface of the fan blades.

In another embodiment according to any of the previous embodiments, an engine case surrounds the gas generator. The engine case includes at least one pylon mount interface for attachment to a pylon mounted underneath a wing.

In another featured embodiment, a gas turbine engine has a propulsor including a fan and a fan drive geared architecture. The fan defines a fan diameter. A gas generator includes a fan drive turbine, which drives the fan through the fan drive geared architecture. The fan drive turbine has a diameter between about 0.35 and about 0.45 the size of the fan diameter.

In another embodiment according to the previous embodiment, the fan drive geared architecture has a gear reduction ratio of greater than about 2.3

In another embodiment according to any of the previous embodiments, the fan drive geared architecture comprises an epicyclic gear box.

In another embodiment according to any of the previous embodiments, a compressor section has at least a first compressor and a second compressor, a combustor in fluid communication with the compressor section, and at least one additional turbine. A first shaft connects the fan drive turbine and the first compressor and a second shaft connects the second compressor and the one additional turbine.

In another embodiment according to any of the previous embodiments, the second shaft rotates at a faster speed than the first shaft.

In another embodiment according to any of the previous embodiments, the fan drive turbine comprises a low pressure turbine and the one additional turbine comprises a high pressure turbine.

In another embodiment according to any of the previous embodiments, the fan drive geared architecture couples the first shaft to the fan at a location upstream of the compressor section.

In another embodiment according to any of the previous embodiments, an engine case surrounds the gas generator. The engine case includes at least one pylon mount interface for attachment to a pylon mounted underneath a wing.

In another embodiment according to any of the previous embodiments, the pylon mount interface comprises at least a front mount beam and a rear mount beam located aft of the front mount beam.

BRIEF DESCRIPTION OF THE DRAWINGS

The disclosure can be further understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein:

FIG. 1 schematically illustrates a geared turbofan engine embodiment.

FIG. 2 schematically illustrates a direct drive turbine engine embodiment.

FIG. 3 shows a side view of a geared turbofan embodiment in one example mounting configuration.

FIG. 4 shows an end view of FIG. 3 in an aft direction looking forward.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates an example gas turbine engine 20 that includes a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28. Alternative engines might include an augmenter section (not shown) among other systems or features. The fan section 22 drives air along a bypass flow path B while the compressor section 24 draws air in along a core flow path C where air is compressed and communicated to a combustor section 26. In the combustor section 26, air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through the turbine section 28 where energy is extracted and utilized to drive the fan section 22 and the compressor section 24.

Although the disclosed non-limiting embodiment depicts a turbofan gas turbine engine, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines; for example a turbine engine including a three-spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive a fan via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that enables a high pressure turbine to drive a high pressure compressor of the compressor section.

The example engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 40 that connects a fan 42 and a low pressure (or first) compressor section 44 to a low pressure (or first) turbine section 46. The inner shaft 40 drives the fan 42 through a speed change device, such as a geared architecture 48, to drive the fan 42 at a lower speed than the low speed spool 30. The high-speed spool 32 includes an outer shaft 50 that interconnects a high pressure (or second) compressor section 52 and a high pressure (or second) turbine section 54. The inner shaft 40 and the outer shaft 50 are concentric and rotate via the bearing systems 38 about the engine central longitudinal axis A.

A combustor 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54. In one example, the high pressure turbine 54 includes at least two stages to provide a double stage high pressure turbine 54. In another example, the high pressure turbine 54 includes only a single stage. As used herein, a “high pressure” compressor or turbine experiences a higher pressure than a corresponding “low pressure” compressor or turbine.

The example low pressure turbine 46 has a pressure ratio that is greater than about 5. The pressure ratio of the example low pressure turbine 46 is measured prior to an inlet of the low pressure turbine 46 as related to the pressure measured at the outlet of the low pressure turbine 46 prior to an exhaust nozzle.

A mid-turbine frame 58 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The mid-turbine frame 58 further supports bearing systems 38 in the turbine section 28 as well as setting airflow entering the low pressure turbine 46.

The core airflow C is compressed by the low pressure compressor 44 then by the high pressure compressor 52 mixed with fuel and ignited in the combustor 56 to produce high speed exhaust gases that are then expanded through the high pressure turbine 54 and low pressure turbine 46. The mid-turbine frame 58 includes vanes 60, which are in the core airflow path and function as an inlet guide vane for the low pressure turbine 46. Utilizing the vane 60 of the mid-turbine frame 58 as the inlet guide vane for low pressure turbine 46 decreases the length of the low pressure turbine 46 without increasing the axial length of the mid-turbine frame 58. Reducing or eliminating the number of vanes in the low pressure turbine 46 shortens the axial length of the turbine section 28. Thus, the compactness of the gas turbine engine 20 is increased and a higher power density may be achieved.

The disclosed gas turbine engine 20 in one example is a high-bypass geared aircraft engine. In a further example, the gas turbine engine 20 includes a bypass ratio greater than about six (6), with an example embodiment being greater than about ten (10). The example geared architecture 48 is an epicyclical gear train, such as a planetary gear system, star gear system or other known gear system, with a gear reduction ratio of greater than about 2.3.

In one disclosed embodiment, the gas turbine engine 20 includes a bypass ratio greater than about ten (10:1) and the fan diameter is significantly larger than an outer diameter of the low pressure compressor 44. It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a geared architecture and that the present disclosure is applicable to other gas turbine engines.

A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section 22 of the engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft., with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFCT’)”—is the industry standard parameter of pound-mass (lbm) of fuel per hour being burned divided by pound-force (lbf) of thrust the engine produces at that minimum point.

“Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.50. In another non-limiting embodiment the low fan pressure ratio is less than about 1.45.

“Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram° R)/(518.7° R)]^(0.5). The “Low corrected fan tip speed”, as disclosed herein according to one non-limiting embodiment, is less than about 1150 ft/second.

The example gas turbine engine includes the fan 42 that comprises in one non-limiting embodiment less than about 26 fan blades. In another non-limiting embodiment, the fan section 22 includes less than about 20 fan blades. Moreover, in one disclosed embodiment the low pressure turbine 46 includes no more than about 6 turbine rotors schematically indicated at 34. In another non-limiting example embodiment the low pressure turbine 46 includes about 3 turbine rotors. A ratio between the number of fan blades 42 and the number of low pressure turbine rotors is between about 3.3 and about 8.6. The example low pressure turbine 46 provides the driving power to rotate the fan section 22 and therefore the relationship between the number of turbine rotors 34 in the low pressure turbine 46 and the number of blades 42 in the fan section 22 disclose an example gas turbine engine 20 with increased power transfer efficiency.

The configuration shown in FIG. 2 is a direct drive turbine engine 25. The direct drive turbine engine 25 includes a fan section 22′, a compressor section 24′, a combustor section 26′, and a turbine section 28′. The fan section 22′ drives air along a bypass flow path B′ while the compressor section 24′ draws air in along a core flow path C′ where air is compressed and communicated to the combustor section 26′. In the combustor section 26′, air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through the turbine section 28′ where energy is extracted and utilized to drive the fan section 22′ and the compressor section 24′.

The direct drive turbine engine 25 generally includes a low speed spool 30′ and a high speed spool 32′ mounted for rotation about an engine central longitudinal axis A′ relative to an engine static structure via several bearing systems 38′. The low speed spool 30′ generally includes an inner shaft that connects a fan 42′ having a plurality of blades and a low pressure (or first) compressor section 44′ to a low pressure (or first) turbine section 46′. The inner shaft or low speed spool 30′ directly drives the fan 42′, that is, the fan 42′ and low pressure turbine section 46′ are driven at the same speed. The high-speed spool 32′ includes an outer shaft that interconnects a high pressure (or second) compressor section 52′ and a high pressure (or second) turbine section 54′. The inner shaft and the outer shaft are concentric and rotate via the bearing systems 38′ about the engine central longitudinal axis A′.

In the direct drive configuration shown in FIG. 2, a fan drive turbine directly drives the fan section 22′, i.e. there is no geared architecture in this configuration. In FIG. 2, the fan drive turbine comprises the low pressure turbine 46′ which is coupled to directly drive the fan 42′.

The geared architecture configuration has increased efficiency that enables the use and fabrication of a smaller low pressure turbine 46 both in diameter and in the number or overall stages as compared to the direct drive turbine engine 25 (FIG. 2), which must rotate at a less efficient speed.

Moreover, the smaller, more efficient low pressure turbine 46 of the geared turbofan engine 20 enables alternate and more efficient mounting configurations. Space limitations for wing mounted engines result from a minimum distance between a bottom of an engine and the runway. Larger landing gear components can be utilized to raise the aircraft and thereby the engine relative to the runway, but larger landing gear components are not a desirable option due to significant weight penalties. Accordingly, as the propulsor fan section 22 grows in size, the mounting options decrease. For engines having the same fan section diameter, the fan drive turbine section of the direct drive engine 25 (FIG. 2) is much larger than the fan drive turbine section of a geared turbofan engine 20 (FIG. 1).

This difference becomes significant when defining a mounting configuration for the engine. The core engine section including the fan drive turbine section can be mounted under the wing, with the fan section extending forward of the wing. The larger fan drive turbine section of a direct drive turbine requires that the engine centerline be spaced a further distance from a bottom surface of the wing as compared to a centerline of a geared turbofan engine with the smaller more efficient fan drive turbine. Even modest reductions in this spacing can enable significant weight savings in smaller landing gear lengths and structures.

The example geared turbofan engine 20 includes a fan diameter 62 (FIG. 1) and an example direct drive engine 25 includes a fan diameter 64 (FIG. 2). In one example configuration, both the fan diameter 62 of the geared turbofan engine 20 and the fan diameter 64 of the direct drive turbine engine 25 are of a common size. Further, in this example, the fan pressure ratio and overall pressure ratio through the core are the same. When these fan diameters 62, 64 and pressure ratios are the same, the geared turbofan engine 20 includes a fan drive turbine diameter 66 (FIG. 1) that is much smaller than a diameter 68 (FIG. 2) of the fan drive turbine for the direct drive engine 25. In one example, for a common fan diameter, the fan drive turbine is about 0.35 to about 0.45 the diameter 62 of the fan 42, wherein a corresponding direct drive engine 25 would include a fan drive turbine between about 0.50 and 0.65 the diameter 64 of the fan 42′.

FIGS. 3-4 show the geared turbofan engine 20 in one example mount configuration. A front mount beam 70 and a rear mount beam 72 are used to connect the engine case 74 to a pylon 76 that is mounted underneath a wing. One relatively important dimension, indicated at 80, is the distance between a bottom surface 82 of the wing and an outermost surface 84 of the fan drive turbine section, that is, low pressure turbine 46. For a fan diameter 64 (FIG. 2) that is the same as the fan diameter 62 for the geared turbofan engine 20 in FIGS. 3-4, the fan drive turbine section, that is, low pressure turbine 46′, would have a comparatively greater size as indicated by an outermost surface 78 of the low pressure turbine 46′. The increased turbine size for the direct drive configuration decreases the wing clearance dimension 80′ when compared to the dimension 80 of the geared turbo fan engine 20.

Thus, the significance of the difference in size of the two different fan drive turbine sections is illustrated with the required spacing of the critical dimension 80′ for a direct drive turbine indicated between the outermost surface 78, shown by the dashed lines, and the bottom surface 82 of the wing. Accordingly, the size of the fan 42′ for a direct drive turbine engine 25 is limited by the size of the fan drive turbine, i.e. the size of the low pressure turbine 46′. As such, the geared turbofan engine 20 with the smaller more efficient fan drive turbine, i.e. low pressure turbine 46, can provide a larger fan in the same space, and/or enable a fan size not possible in a direct drive gas turbine engine 25.

Although an example embodiment has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of the claims. For that reason, the following claims should be studied to determine their true scope and content. 

What is claimed is:
 1. A gas turbine engine comprising: a fan configured to deliver airflow to a bypass passage; a core engine configured to rotate the fan, the core engine including: a high pressure turbine section configured to drive a high pressure compressor section, and a low pressure turbine section configured to drive the fan and a low pressure compressor section, the fan and the low pressure turbine section configured to rotate at a common speed and in a common direction; and wherein the gas turbine engine has a bypass ratio of greater than about 10, the fan has a fan diameter, Dfan, the low pressure turbine section has a turbine diameter, Dturb, and the fan diameter Dfan and the turbine diameter Dturb have an interdependence represented by a scalable ratio Dturb/Dfan that is between 0.5 and 0.65.
 2. The gas turbine engine as recited in claim 1, wherein the turbine diameter Dturb is defined by an outer case surface of the low pressure turbine section.
 3. The gas turbine engine as recited in claim 1, wherein the fan includes a plurality of fan blades, and the fan diameter Dfan is defined by outer peripheral surfaces of the fan blades.
 4. The gas turbine engine as recited in claim 1, wherein the high pressure turbine section includes two stages.
 5. The gas turbine engine as recited in claim 1, wherein the low pressure compressor section includes a greater number of stages than the high pressure turbine section, and includes fewer stages than the high pressure compressor section.
 6. The gas turbine engine as recited in claim 1, wherein the low pressure turbine section has a greater number of stages than the low pressure compressor section.
 7. The gas turbine engine as recited in claim 6, wherein the fan has a pressure ratio of less than about 1.45.
 8. The gas turbine engine as recited in claim 7, wherein the fan includes fewer than 20 fan blades.
 9. The gas turbine engine as recited in claim 1, wherein the low pressure turbine section has a pressure ratio of greater than about
 5. 10. A gas turbine engine comprising: a fan having fewer than 20 fan blades situated at an inlet of a bypass passage; a core engine configured to rotate the fan, the core engine including: a high pressure turbine section configured to drive a high pressure compressor section, and a low pressure turbine section configured to drive the fan and a low pressure compressor section, the low pressure turbine section having a greater number of stages than the low pressure compressor section; and wherein the fan has a fan diameter, Dfan, the low pressure turbine section has a turbine diameter, Dturb, and the fan diameter Dfan and the turbine diameter Dturb have an interdependence represented by a scalable ratio Dturb/Dfan that is between 0.5 and 0.65.
 11. The gas turbine engine as recited in claim 10, wherein the turbine diameter Dturb is defined by an outer case surface of the low pressure turbine section.
 12. The gas turbine engine as recited in claim 10, wherein the fan diameter Dfan is defined by outer peripheral surfaces of the fan blades.
 13. The gas turbine engine as recited in claim 10, wherein the high pressure turbine section includes two stages.
 14. The gas turbine engine as recited in claim 10, wherein the low pressure compressor section includes a greater number of stages than the high pressure turbine section, and includes fewer stages than the high pressure compressor section.
 15. The gas turbine engine as recited in claim 14, wherein the fan has a pressure ratio of less than about 1.5.
 16. The gas turbine engine according to claim 10, wherein the gas turbine engine has a bypass ratio of greater than about
 10. 17. The gas turbine engine according to claim 16, wherein the low pressure turbine has a pressure ratio of greater than about
 5. 18. A method of designing a gas turbine engine comprising: providing a fan configured to deliver airflow to a bypass duct; providing a core engine configured to rotate the fan, the core engine including: a high pressure turbine section configured to drive a high pressure compressor section, and a low pressure turbine section configured to rotate the fan at a common speed and in a common direction; wherein the fan has a fan diameter, Dfan, the low pressure turbine section has a diameter, Dturb, and the fan diameter Dfan, and the low turbine section diameter Dturb have an interdependence represented by a scalable ratio Dturb/Dfan that is between 0.5 and 0.65; and wherein the fan is configured to deliver a portion of air into the core engine, and a portion of air into the bypass duct, and a bypass ratio, which is defined as a volume of air passing to the bypass duct compared to a volume of air passing into the core engine, is greater than
 10. 19. The method as recited in claim 18, wherein the fan has a pressure ratio of less than about 1.5 and the low pressure turbine has a pressure ratio of greater than about
 5. 20. The method as recited in claim 18, comprising a low pressure compressor section driven by the low pressure turbine section, wherein the low pressure compressor section includes a greater number of stages than the high pressure turbine section, and includes fewer stages than the high pressure compressor section, and the high pressure turbine section includes two stages. 